1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially in an industrial gas turbine engine, compressed air is delivered to a combustor and burned with a fuel to produce an extremely hot gas flow. The hot gas flow is passed through a multiple stage turbine to extract mechanical energy. The engine efficiency can be increased by increasing the temperature of the hot gas flow entering the turbine. One of the major problems with the design of gas turbine engines is forming the first stage stator vanes and rotor blades from materials that can withstand the extreme high temperature of the hot gas flow. In order to overcome the limitations due to the material properties, complex internal cooling circuits have been proposed to provide high levels of cooling for these airfoils while minimizing the amount of cooling air used. Since the pressurized cooling air is typically diverted from the compressor of the engine, which is compressed air that is not used to perform work, using less air from the compressor for cooling will also increase the engine efficiency.
Prior Art turbine airfoils near wall cooling utilized in an airfoil main body is constructed with radial flow channels plus re-supply holes in conjunction with film discharge cooling holes. As a result of this cooling construction approach, span-wise and chord-wise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, single radial channel flow is not the best method of utilizing cooling air, resulting in a low convective cooling effectiveness. U.S. Pat. No. 5,660,524 issued to Lee et al on Aug. 26, 1997 and entitled AIRFOIL BLADE HAVING A SERPENTINE COOLING CIRCUIT AND IMPINGEMENT COOLING discloses a turbine airfoil blade with generally longitudinally extending coolant passageways (#40, 42, and 44 in this patent) with first and second impingement chambers (#53 and 60 in this patent) located on the pressure side and the suction side of the blade adjacent to the coolant passageway. The two impingement chambers also extend along the entire span-wise direction of the blade from the root to the blade tip. One problem with this design is that the blade may have hot spots along the span-wise direction. Because the impingement chamber is one long passage, some areas of the blade along the span-wise direction may be under-cooled while others may be over-cooled.
U.S. Pat. No. 6,773,230 B2 issued to Bather et al on Aug. 10, 2004 and entitled AIR COOLED AEROFOIL discloses a turbine airfoil with a central cooling air supply channel and a series of cooling wall cavities spaced along the airfoil wall and connected to the cooling air supply channel by impingement holes. The impingement cavities can be separated into a plurality of compartments spaced along the airfoil span-wise direction in order to increase the efficiency of such a cooling arrangement (see column 3, line 42 of this patent). In the Bather et al patent, the source of cooling air supply is only connected to the central cavity (#34 in this patent), and this central cavity is in direct fluid communication with the film cooling holes that provide cooling for the leading edge showerhead arrangement. Also, the impingement cooling air passes into the second cavity (#26 in this patent) which is located downstream from the first or supply cooling air cavity. Therefore, a series flow is formed that passes from the first cooling air supply cavity 34, into the impingement cavities 24 and 28, into the second cavity 26, and then into a trailing edge cavity 26 and out through exit cooling holes 44 in the trailing edge of the airfoil. This is a long flow path for the cooling air, which results in lower efficiency because the cooling air heats up before reaching the middle and trailing edge portions of the airfoil.
It is an object of the present invention to provide a turbine airfoil with a near wall cooling arrangement for a turbine airfoil main body region that will greatly reduce the airfoil main body metal temperature and thus reduce the cooling flow requirement and improve the turbine efficiency.
It is another object of the present invention to provide for a turbine airfoil in which the airfoil is cooled by a cooling air circuit that uses convection in series with impingement cooling and film cooling to maximize the heat transfer coefficient while minimizing the amount of cooling air used.